(1) Field of the Invention
The present invention relates to a system and to a method for controlling pitching stabilizer means, to a stabilizer assembly, and to an aircraft fitted with the system. More particularly, the invention lies in the narrow technical field of means for stabilizing rotorcraft in pitching.
(2) Description of Related Art
By way of example, a rotorcraft conventionally comprises a fuselage that extends longitudinally from a front end to a tail end on either side of an anteroposterior plane of symmetry, and extends in a vertical direction from a bottom portion fitted with landing gear to a top portion fitted with a rotary wing. The rotary wing may include a main rotor providing lift and propulsion.
A rotorcraft may also include a tail rotor at its tail end. The tail rotor serves in particular to counter the yaw torque exerted by the main rotor on the fuselage. The tail rotor also makes it possible to control yaw movements of the helicopter.
The rotary wing and the tail rotor can be controlled by flight controls connected by mechanical linkages to the rotary wing and/or to the tail rotor. Each mechanical linkage may be associated with an automatic control system including an actuator referred to as a “series” actuator and an actuator referred to as a “parallel” actuator or indeed as a “trim” actuator.
The series actuator is controlled by a computer to stabilize the aircraft. Such a series actuator has limited authority, possessing a stroke that does not enable it to cover the entire stroke of the associated flight control. Nevertheless, this limited authority remains sufficient to stabilize the aircraft about a required central position.
The series actuator is also a fast-moving actuator so as to enable the mechanical linkage to be adjusted quickly in order to stabilize the aircraft, e.g. following a gust of wind.
The series actuator is arranged in series in the mechanical linkage so as to be transparent for a pilot. In other words, the movements of a series actuator tend not to be felt by the pilot via the flight controls.
In contrast, the trim actuator can take the place of a pilot to act on the mechanical linkage over the full stroke of the associated flight control. Thus, a trim actuator has full authority over the movement of the mechanical linkage. A pilot can feel the action of the trim actuator. Since the trim actuator is arranged in parallel with the mechanical linkage, the trim actuator moves at least one flight control by means of the mechanical linkage on which the trim actuator acts.
A computer may then control the trim actuator to take the place of a pilot. Nevertheless, the computer may also call on the trim actuator to re-center the mechanical linkage as a function of movements requested by the series actuator, in particular when the series actuator reaches an abutment.
Conventionally, the trim actuator is a slow-moving actuator, as contrasted to the fast nature of the series actuator.
A rotorcraft sometimes also has additional stabilizer surfaces. For example, it is common practice to fit a rotorcraft with a stabilizer surface for stabilizing yaw movements.
Such a yaw movement stabilizer surface is generally referred to as a “fin”.
Likewise, a rotorcraft sometimes has means for balancing and stabilizing pitching movements, referred to merely as “stabilizer means”. Stabilizer means may present an angle of absolute value lying in the range 0 degree (°) plus or minus 90° relative to said anteroposterior plane. The pitching stabilizer means may optionally comprise two pitching stabilizer surfaces extending symmetrically on either side of an anteroposterior plane of symmetry of the rotorcraft and being orthogonal to the anteroposterior plane, or else possibly presenting a V-shape, for example.
Such stabilizer means for stabilizing pitching movements are sometimes referred to as a “horizontal stabilizer”, or more simply merely as a “stabilizer”. The term “stabilizer” is in widespread use since the stabilizer means are not necessarily horizontal. It is also possible to use the term “pitching stabilizer means”.
A stabilizer may comprise at least one airfoil surface passing through the rear end of the aircraft in a transverse direction, or indeed at least one non-through airfoil surface extending transversely from said rear end on one side only of said anteroposterior plane.
In hovering flight, balancing the pitching moment of the rotorcraft relative to a fixed point involves two major components: the moment due to the weight of the aircraft and the moment due to the main propulsion and lift rotor, which, for a given weight of rotorcraft, is proportional to the angle of tilt of the main rotor relative to the vertical direction. Furthermore, variations in the position of the center of gravity of the helicopter give rise to variations in the attitude of the aircraft.
While the rotorcraft is in cruising flight, another component of the pitching moment is involved: the aerodynamic moment exerted by the airframe of the aircraft. As a result of varying the angle of incidence of the airframe relative to an upstream flow of air, the aerodynamic pitching moment tends to cause the aircraft to depart from its equilibrium position. This unstable component has the effect of increasing variations in the longitudinal attitude that are associated with the position of the center of gravity, as compared with the variations that are observed while hovering.
These variations of attitude have negative consequences. Excessive nose-down attitudes increase the aerodynamic drag of the rotorcraft, and consequently reduce its maximum speed. They also give rise to an impression of discomfort for the crew and for passengers. Excessive nose-up attitudes give rise to large moments on the mast and on the hub of the main rotor, with unfavorable consequences on the lifetimes of those elements.
The pitching stabilizer means located towards the rear of the aircraft seek to compensate instability in the pitching moment of the fuselage and to keep the attitude of the rotorcraft in equilibrium.
Stabilizer means are complex to design. In order to optimize the performance of an aircraft at high speed, the pitching stabilizer means are designed so as to obtain a longitudinal attitude that is close to a level attitude. Nevertheless, such a design can be penalizing for the operation of the main rotor, it being understood that it is more desirable to have a large nose-down attitude.
Furthermore, the design must be satisfactory for various configurations of weight, altitude, position of center of gravity, and possibly potential external aerodynamic configurations of the aircraft.
The effectiveness of the pitching stabilizer means may possibly be maximized by increasing wing area so as to reduce the effects of disturbances associated with variations in the weight and the position of the center of gravity of the rotorcraft.
Nevertheless, such a solution is limited, e.g. because of the “attitude hump” phenomenon known to the person skilled in the art, that results from interactions between the main rotor and the pitching stabilizer means. Furthermore, a large wing area tends to give rise to large variations in the attitude of the aircraft while the aircraft is climbing or descending.
Manufacturers have sought to remedy those drawbacks by creating a device for controlling pitching stabilizer means in such a manner as to balance a rotorcraft in pitching, while simultaneously keeping control over its performance and the loads applied to the hub of a main rotor, while being unaffected by variations in the position of the center of gravity, and while complying with the above-mentioned constraints needing the size of said airfoil surface to be limited.
Thus, Documents U.S. Pat. No. 2,424,882 and GB 657 796 provide for a lever that is mechanically connected to pitching stabilizer means in order to control the angle of incidence of the stabilizer means.
Those documents suggest manual piloting. Nevertheless, the complexity of modern rotorcraft can lead to the angle of incidence of the pitching stabilizer means being servo-controlled depending on flight conditions.
Document FR 2 456 663 provides for using an actuator device to servo-control the angle of incidence of a chord plane. That actuator device is provided with two motors to cause a rod to move in translation.
Certain aircraft make use of two electric actuators mounted back to back, one being fastened to the tail boom of said aircraft and the other being fastened to pitching stabilizer means.
Actuator redundancy makes it possible to operate the stabilizer in the event of one of the actuators failing.
The technological background remote from the invention also includes the following documents: FR 2 603 866, US 2009/0206197, FR 2 809 372, U.S. Pat. No. 4,834,319, WO 2008/038037, and U.S. Pat. No. 6,461,265.
Document FR 2 603 866 relates to a system for controlling airplane control surfaces in a context of so-called “fly-by-wire” assisted piloting associated with an auxiliary control that is mechanical.
It should be observed that the technical field of airplanes is remote from the technical field of rotorcraft. In particular, a rotorcraft responds to flight control orders much more quickly and strongly than does an airplane. Furthermore, an airplane possesses natural stability, as contrasted to the unstable nature of rotorcraft, which tends to require stability to be controlled continuously.
Document US 2009/0206197 describes two actuators connected to a flap, the two actuators being connected to each other.
Document FR 2 809 372 relates to an aircraft with electric flight controls provided with an actuator for moving a control surface.
Document U.S. Pat. No. 4,834,319 describes an actuator having a first motor with authority over a moving surface that is limited by an abutment, and a second motor for adjusting the margin of the first motor relative to the abutment.